Low shock strength propulsion system

ABSTRACT

A supersonic nacelle design employing a bypass flow path internal to the nacelle and around the engine is disclosed herein. A set of aerodynamic vanes may be used to shape a supersonic airflow within a bypass around an engine. The vanes may be capable of compressing the supersonic airflow into a subsonic airflow, direct the subsonic airflow around the engine, and then expand the subsonic airflow into a supersonic exhaust. The vanes may shape the airflow by reducing sonic boom strength, cowl drag, and airframe interference drag.

This application is a Continuation of U.S. Non-Provisional patentapplication Ser. No. 12/605,071, filed Oct. 23, 2009, and entitled “LowShock Strength Propulsion System” which is a Continuation-in-Part ofU.S. Non-Provisional patent application Ser. No. 12/000,066, filed Dec.7, 2007 and entitled “Low Shock Strength Inlet,” and U.S.Non-Provisional patent application Ser. No. 12/257,982, filed Oct. 24,2008 and entitled “Low Shock Strength Propulsion System,” both of whichclaim priority to U.S. Provisional Patent Application 60/960,986, filedOct. 24, 2007 and entitled “Supersonic Nacelle.” These applications arehereby incorporated by reference in their entirety and are commonlyowned by the assignee.

FIELD OF THE INVENTION

Embodiments of the invention are related to supersonic inlets andnozzles for supersonic engines and more particularly to supersonicinlets and nozzles configured with a bypass for reduced sonic boomstrength.

BACKGROUND OF THE INVENTION

Many supersonic aircraft employ gas turbine engines that are capable ofpropelling the aircraft at supersonic speeds. These gas turbine engines,however, generally operate on subsonic flow in a range of about Mach 0.3to 0.6 at the upstream face of the engine. In supersonic applications, anacelle is used to encompass the engine and incorporates an inlet and anozzle. The inlet decelerates the incoming airflow to a speed compatiblewith the requirements of the gas turbine engine. To accomplish this, asupersonic inlet is comprised of a compression surface and correspondingflow path, used to decelerate the supersonic flow into a strong terminalshock. Downstream of the terminal shock, subsonic flow is furtherdecelerated using a subsonic diffuser to a speed corresponding with thein-flow requirements of the gas turbine engine. The exhaust from theengine is then accelerated again using the nozzle.

Traditional supersonic propulsion system design methods minimize thediameter and structural weight of the nacelle while maximizing grossthrust. In doing so, the amount of flow captured by the inlet is limitedto only that demanded by the engine with an additional small amount fornacelle purge and cooling. A measurement of inlet operation efficiencyis the total pressure lost in the air stream between the entrance sideand the discharge side of the inlet. The total pressure recovery of aninlet is defined by a ratio of the total pressure at the discharge tothe total pressure at the free stream. Maximizing the total pressurerecovery leads to maximizing gross engine thrust, thus improving theperformance of the propulsion system

FIG. 1 schematically illustrates a cross-sectional view of a traditionalnacelle 10, having an external compression inlet 11 and nozzle 14surrounding an engine 16. The external compression inlet 11 compressesand decelerates the supersonic flow to the face of the engine 16. Theinlet 11 includes the leading edge 12 of the compression surface and thecowl 13 forming the inlet opening of the inlet 11. The output of theengine 16 is then accelerated by the nozzle 14, creating the necessarythrust to propel the aircraft at supersonic speeds. The nacelle 10 isoften designed to cover around the protruding engine parts 18, which mayinclude engine components such as gear boxes and other components knownto those of skill in the art. As shown in FIG. 1, the engine 16 may be aconventional turbofan type engine featuring approximately 15,000 lb_(f)of maximum takeoff thrust and a moderate fan-to-compressor flow ratio of3.

Unfortunately, the traditional nacelle design for a supersonic engineconfiguration often generates strong shocks off the supersonic inlet andfrom the body of the nacelle. A traditional approach to supersonic inletdesign typically employs shock-on-lip focusing. As would be understoodby those of skill in the art, shock-on-lip focusing involves designing acompression surface configuration of an external compression inlet suchthat the inlet-generated shocks (that occur at a supersonic designcruise speed) meet at a location immediately forward of the cowlhighlight or the cowl lip.

FIG. 2 schematically illustrates the bottom half of the cross-sectionalview in FIG. 1 and how shock waves and expansion regions are generatedby the nacelle 10 and how the shape of the nacelle 10 may generateadditional shock waves and expansion regions. As is understood by thoseof skill in the art, these shock waves can coalesce into a strongershock wave as the shock waves propagate away from the aircraft duringsupersonic flight. These shock waves can also propagate into theaircraft surface, creating localized regions of interference drag. Theshock wave 20 is generated by the leading edge 12 of the initialcompression surface of the inlet 11. The wave 20 may coalesce with theshock wave 22, generated by the cowl 13, and potentially the shock waves24 and 26. These shock waves may then coalesce with shock wavespropagating from the airframe itself, eventually creating a sonic boomas heard at ground level.

The shock wave 22 is often referred to as the cowl shock, the strengthof which may be directly related to the cowling angle A. In addition,any increase in cowling angle results in additional inlet frontal area,which increases inlet drag as speed increases. This adverse trend is akey reason why conventional external compression inlets lose viabilityat high supersonic Mach numbers. Other shock waves, such as shock wave24, and expansion regions, identified in region 25, are often caused bychanges in the shape and diameter of the nacelle 10, especially as thenacelle attempts to cover the protruding engine parts 18. The shock wave26 is generated off the trailing edge 15 of the nozzle. As is understoodby those of skill in the art, the strength of this shock wave 26 isproportional to the nozzle cowling angle B, often referred to as thenozzle boat tail angle.

Unfortunately, these shock waves combine with those from the airframe tocreate a louder overall sonic boom signature and more interference dragbetween the nacelle and the remainder of the vehicle. The stronger theshock waves, the more difficult they become to control and attenuate andthe more likely they are to produce additional drag and sonic boomnoise.

One way to control drag, as discussed in U.S. Pat. No. 6,793,175 toSanders, involves configuring the inlet to minimize the shape and sizeof the cowl. The configuration of the inlet initially resembles acircumferential sector of an axisymmetric intake, but switches thelocation of compression surface to the outer radius and disposes thecowling on the inner radius in a higher performance, 3-D geometry. Thefact that the cowl is located on the inner radius reduces the physicalarc of the cowl. Problems with this method include the aircraftintegration challenges created by the 3-D geometry, such as the factthat the cross-sectional shape may be more difficult to integrate from apackaging perspective compared to an equivalent axisymmetric design forpodded propulsion systems. In addition, the complex inlet shape islikely to create complex distortion patterns that require either largescale mitigating techniques in the subsonic diffuser or the use ofengines with more robust operability characteristics.

Another way to control drag by reducing the cowl lip angle is based ondecreasing the flow turn angle by increasing the inlet terminal shockMach number. The improvement in drag reduction is often negated by thereduction in pressure recovery resulting from the stronger terminalshock. In addition, increasing the terminal shock Mach number at thebase of the shock also encounters significant limitations in practicedue to viscous flow effects. Higher terminal shock Mach numbers at thebase of the shock aggravate the shock-boundary layer interaction andreduce shock base boundary layer health. The increase in shock strengthin the base region also reduces inlet buzz margin, reducing subcriticalflow throttling capability. Additionally, the increase in terminal shockMach number will most likely require complex boundary layer managementor a complex inlet control system.

Inlet compression surfaces are typically grouped into two types:straight or isentropic. A straight surface has a flat ramp or conicsections that produce discrete oblique or conic shocks, while anisentropic surface has a continuously curved surface that produces acontinuum of infinitesimally weak shocklets during the compressionprocess. Theoretically, a traditional isentropic compression surface canhave better pressure recovery than a straight surface designed to thesame operating conditions, but real viscous effects can reduce theoverall performance of the isentropic surface inlets and result inpoorer boundary layer health.

FIG. 3 schematically represents a perspective view of a enginearrangement 30 representative of a high specific thrust militaryturbofan engine of approximately 11,000 lb_(f) maximum takeoff thrustclass (non-afterburning). The arrangement 30 may include a nacelle 32,having a traditional inlet 34 and nozzle 36. As can be seen from FIG. 3,the nacelle must be configured to encompass the protruding parts 40 ofthe engine 38. Moreover, the non-optimal matching between the intakearea and the maximum nacelle diameter creates a large forward cowlingprofile that results in high drag and strong shock generation. Likewise,the non-optimal matching between exhaust area and maximum nacellediameter causes a large nozzle boat tail angle, resulting in high dragand strong expansion and re-shock.

FIG. 4 schematically illustrates a perspective view of the engine 38from FIG. 3 installed on the vertical stabilizer of a supersonicaircraft 42. The nacelle 32 is configured to encompass the protrudingengine parts 40, creating a generally asymmetric configuration, which asdiscussed above may contribute to the generation of shock waves and thestrength of a resulting sonic boom. While such performance may beacceptable for military aircraft or other such applications, thegeneration of strong sonic booms in the civil aviation arena isundesirable.

SUMMARY OF THE INVENTION

Embodiments of the invention may include a nacelle configuration thatemploys a bypass configured to capture, route, and exhaust the largeamount of excess airflow within and through an aircraft nacelle, butexternal to the engine. The inclusion of a bypass stream enables nacelleshape tailoring that would otherwise not be possible for a propulsionsystem employing a conventional, single flow stream design. Whendesigning a supersonic nacelle, embodiments of the invention capitalizeon a broadened trade space that considers sonic boom impact, cowl drag,airframe interference drag, subsystem complexity, and structural designtechniques.

In one embodiment of the invention, a supersonic nozzle for a supersonicengine comprises: an outer wall, a bypass wall disposed within the outerwall, a set of struts configured to couple the outer wall with thebypass wall. The bypass wall may be configured to separate an airflowinto a primary flow portion and a bypass flow portion, such that theprimary flow portion passes through the supersonic engine and the bypassflow portion passes through a bypass. The set of struts also may beconfigured to tailor a direction of the bypass flow portion.

In another embodiment of the invention, a low shock supersonic nacellecomprises: an engine, an outer wall, a bypass wall disposed within theouter wall, a set of struts configured to couple the outer wall with thebypass wall, an inlet defined front portions of the outer wall and thebypass wall, and a nozzle defined by rear portions of the outer wall andthe bypass wall. The inlet may be configured to decelerate an incomingairflow to a speed compatible with the engine, while the nozzle may beconfigured to accelerate an exhaust from the engine and a bypass. Thebypass wall may be configured to divide the incoming airflow into aprimary flow portion directed into the engine and a bypass flow portiondirected into the bypass.

In another embodiment of the invention, a method for decelerating asupersonic flow for a supersonic propulsion system comprises: cruisingat a predetermined supersonic speed, receiving a supersonic flow in aninlet, splitting a subsonic flow into a primary flow portion and abypass flow portion, diffusing the primary flow portion with a diffuserto a predetermined speed suitable for an engine, expanding the primaryflow portion after the primary flow portion leaves the engine andreaches a nozzle, and directing the bypass flow portion into asubstantially circumferentially uniform pattern prior to exhaust. Theinlet may have a compression surface, a bypass splitter, and a cowl lipthat is spatially separated from the compression surface. The bypassflow portion may receive and capture a substantial region of flowdistortion created by the inlet.

In another embodiment of the invention, a nacelle for a supersonicengine comprises: an outer wall defining a closed volume, a bypass walldisposed within the volume of the outer wall, and a set of vanesconfigured to couple the outer wall and the bypass wall. The set ofvanes defines a bypass between the outer wall and the bypass wall and isconfigured to compress a bypass airflow into a subsonic flow.

In another embodiment of the invention, a low shock supersonic nacellecomprises: an engine, an outer wall defining a closed volume, a bypasswall disposed within the volume of the outer wall and configured tosupport the engine, a set of vanes configured to couple the outer wallwith the bypass wall, an inlet, and a nozzle.

In another embodiment of the invention, a method of decelerating asupersonic flow for a supersonic propulsion system comprises: cruisingat a predetermined supersonic speed, splitting a supersonic inlet flowinto a primary flow portion and a bypass flow portion, diffusing theprimary flow portion with a diffuser to a predetermined speed suitablefor an engine, compressing the bypass flow portion into a subsonicbypass flow, expanding the primary flow portion after the primary flowportion leaves the engine and reaches a nozzle, directing the subsonicbypass flow around the engine, and expanding the subsonic bypass flow toa supersonic bypass exhaust.

BRIEF DESCRIPTION OF THE DRAWINGS

The patent or application file contains at least one drawing executed incolor. Copies of this patent or patent application publication withcolor drawing(s) will be provided by the Office upon request and paymentof the necessary fee.

It is believed that embodiments of the invention will be betterunderstood from the following description taken in conjunction with theaccompanying drawings, which illustrate, in a non-limiting fashion, thebest mode presently contemplated for carrying out embodiments of theinvention, and in which like reference numerals designate like partsthroughout the Figures, wherein:

FIG. 1 schematically illustrates a cross-sectional view of a traditionalnacelle;

FIG. 2 schematically illustrates another cross-sectional view of atraditional nacelle with shock waves and expansion regions propagatingoff the nacelle;

FIG. 3 schematically illustrates a perspective view of an enginearrangement and traditional nacelle design;

FIG. 4 schematically illustrates a perspective view of the engine andnacelle from FIG. 3 installed on a supersonic aircraft;

FIG. 5 schematically illustrates a cross-sectional view of a nacelle inaccordance with an embodiment of the invention;

FIG. 6 shows the nacelle from FIG. 5 overlaid by the traditional nacellefrom FIG. 1;

FIG. 7A schematically illustrates a cross-section of an inlet inaccordance with embodiments of the invention;

FIG. 7B schematically illustrates a side elevation view of a supersonicaircraft inlet entrance;

FIG. 7C schematically illustrates a side view of an inlet in accordancewith an embodiment of the invention;

FIG. 8 illustrates a plane view of the inlet in FIG. 7A from the frontof the inlet looking in the aft direction;

FIGS. 9A and 9B show perspective views of the inlet of FIG. 7A with theout wall of the nacelle drawn transparent;

FIG. 9C illustrates another plane view of the inlet in FIGS. 7A, 8, and9A-B with the engine face shown at the back of the inlet;

FIG. 10A illustrates a Computational Fluid Dynamics (CFD) solution forthe inlet in FIG. 7A at a static or low speed condition;

FIG. 10B shows a graph plotting the axial component of the velocity ofthe flow within the diffuser shown in FIG. 10A against the radialdistance from the center of the nacelle;

FIG. 11A illustrates a CFD solution for the inlet in FIG. 7A at asupersonic design speed of Mach 1.7;

FIG. 11B shows a graph plotting the axial component of the velocity ofthe flow within the diffuser shown in FIG. 11A against the radialdistance from the center of the nacelle;

FIG. 11C illustrates a Mach color CFD solution of an inlet with aconventional cowl;

FIG. 11D illustrates a Mach color CFD solution of an externalcompression inlet with a zero-angle cowl in accordance with anembodiment of the invention;

FIG. 12 schematically illustrates a cross-sectional view of a nozzle 100with a bypass path in accordance with an embodiment of the invention;

FIG. 13 schematically illustrates a plane view of the nozzle 100 fromaft of the nozzle looking forward or upstream;

FIGS. 14A and 14B show perspective views of the nozzle 100 from FIGS. 12and 13 with the outer surface of the nozzle drawn as transparent;

FIG. 15 illustrates a CFD solution at freestream speed of Mach 1.7 ofthe internal flowpath and external flow region for a representativetraditional nozzle design featuring a large nozzle boat tail angle;

FIG. 16 illustrates a CFD solution at freestream speed of Mach 1.7 ofthe internal flowpath and external flow region for the nozzle shown inFIG. 12;

FIG. 17 illustrates a CFD solution for the flow around bypass struts 112shown in FIGS. 12, 13, and 14A-B;

FIG. 18 schematically illustrates a visually exaggerated 2 degree radialsection of the nozzle shown in FIGS. 12, 13, and 14A-B;

FIG. 19 illustrates a cylindrical coordinate view of a nacelle bypassgeometry in accordance with another embodiment of the invention;

FIG. 20 illustrates an isometric front view of the nacelle bypassgeometry of FIG. 19;

FIG. 21 illustrates an isometric rear view of the nacelle bypassgeometry of FIG. 19;

FIG. 22 illustrates a side view of the nacelle bypass geometry of FIG.19;

FIG. 23 illustrates a front view of an inlet section of the nacellebypass geometry of FIG. 19;

FIG. 24 illustrates a rear view of a nozzle section of the nacellebypass geometry of FIG. 19;

FIG. 25 illustrates a CFD solution for a nacelle geometry as shown inFIGS. 19-24;

FIG. 26 illustrates a CFD solution for a nacelle geometry as shown inFIGS. 19-24; and;

FIG. 27 illustrates a CFD solution for an inlet of a nacelle geometry asshown in FIGS. 19-20 and 22-23.

DETAILED DESCRIPTION OF THE INVENTION

The present disclosure will now be described more fully with referenceto Figures in which various embodiments of the invention are shown. Thesubject matter of this disclosure may, however, be embodied in manydifferent forms and should not be construed as being limited to theembodiments set forth herein.

Embodiments of the invention relates to supersonic nacelle and engineconfigurations that include a bypass flow around the engine. Designconsiderations may be used when employing a nacelle incorporating abypass flow around the engine. For example, an expanded design space mayinclude sonic boom impact, cowl drag, airframe interference drag,subsystem complexity, and alternative structural design techniques, allof which may be optimized against a propulsion system configuration thatfavors a more streamlined, albeit enlarged, nacelle shape. Growing theforward cowling diameter to better streamline the forward nacelleresults in additional captured airflow that cannot be used by theengine. Without a system to efficiently dispose of this additionalcaptured airflow, in the known prior art, the excess flow spills aroundthe exterior of the cowl lip, creating higher drag and defeating theobjective of a lower sonic boom signature. To avoid thesespillage-related issues, in embodiments of the invention, the additionalflow is efficiently routed internally through the nacelle and around theengine, eventually exhausting back to freestream.

As would be understood by those of skill in the art, the additional aircaptured by the inlet does not pass through the turbomachinery and,consequently, is not energized. Losses along the internal flowpathprevent the complete re-expansion of the flow back to freestreamsupersonic speed upon exit from the nozzle. These losses createadditional drag. However, this additional ram drag, along with theincreased skin fiction that results from the larger nacelle surfacearea, trades against the reduction in cowl drag and airframeinterference drag in addition to a potential increase in engine thrustresulting from improvement in the total pressure recovery in the primaryflowpath. A properly designed bypass system may be used to minimize oreliminate any performance penalty while significantly reducing thecontribution of the propulsion system to the overall

In accordance with embodiments of the invention, a bypass may beconfigured to enhance the ability to shape or tailor the outer surfaceof the nacelle for improved sonic boom characteristics. As a result,embodiments of the invention may include a more streamlined(‘stovepipe’) nacelle profile that may provide improved matching betweenmaximum nacelle cross-sectional area and the cross-sectional areas ofengine intake and exhaust. Improved area matching reduces the localsloping of the nacelle outer surface which produces a reduction incompression (shock) and expansion region strength. Referring back toFIG. 1, the cross-sectional area of the intake defined by the cowl 13does not match well to the maximum nacelle diameter because thecross-sectional area of the nacelle in the region of the engine mustaccommodate the overwhelming volume requirements of the external enginehardware, especially the gearbox. Therefore, the nacelle 10 growsdramatically in diameter around the engine 16 in order to encompass theprotruding engine parts 18. The nozzle 14 is sized to pass the engineexhaust at exit velocities necessary to meet performance and sonic boomrequirements: the exhaust flow is usually fully expanded at designcondition to maximize thrust and to minimize the exhaust stream'sdisturbance of the external flow field. As shown in FIG. 1, the nozzle14 dramatically reduces in diameter until the trailing edge 15, whichdefines the exit cross-sectional area of the nozzle 14.

Contrary to the traditional nacelle design shown in FIG. 1, nacelles inaccordance with embodiments of the invention may be streamlined suchthat they produce weaker shocks and expansion zones at supersonicspeeds. The nacelle design may also be configured to produce lessnacelle pressure drag and airframe interference drag, when compared to aconventional nacelle that bulges outward to be form-fitted aroundprotruding engine parts, such as the gearbox and other bulky hardwaremounted to the exterior of the engine.

In accordance with embodiments of the invention, the outer surface ofthe nacelle may be configured to encompass the entire engine, includingthose parts that would traditionally create protrusions on the nacelle.FIG. 5 schematically illustrates a cross-sectional view of a nacelle inaccordance with an embodiment of the invention which encloses an engine52. The bypass flow fraction for the embodiment shown in FIG. 5 may belarge, with about one part of flow captured by a bypass 58 for every twoparts ingested by the engine, giving a bypass percentage of about 50percent. The bypass is a portion of the flowpath internal to the nacellethat does not direct flow into, through, or out of the engine. Theengine 52 shown in FIG. 5 is the same as that shown in FIG. 1. It shouldbe understood that embodiments of the invention may be applied to anyair-breathing propulsion system configured for supersonic flight. Thesepropulsion systems could employ conventional turbojet and turbofanengines, combined cycle engines, or ramjets. Propulsion system employingvariable cycle engines that use variable fan blade tip geometry may alsobe used.

The nacelle 50 shown in FIG. 5 also includes an inlet module 54 and anozzle module 56. In accordance with embodiments of the invention, thebypass 58 may be configured to bypass flow around the engine 52 from theinlet 54 to the nozzle 56. The bypass 58 allows the overall design ofthe nacelle 50 to be more cylindrical from cowl 60 on the inlet 54 tothe trailing edge 62 on the nozzle 56.

The arrangement shown in FIG. 5, which may be configured to approximatea straight pipe configuration, has dramatically less variability in thediameter of the nacelle from the intake area defined by the cowl 60 tothe exhaust area defined by the nozzle trailing edge 62. FIG. 6illustrates this difference by overlaying the traditional nacelle design10 from FIG. 1 on top of the nacelle 50 from FIG. 5. As is clear fromFIG. 6, the nacelle 50 is generally larger in diameter but is a muchmore streamlined design, exhibiting less change in outside circumferenceof the nacelle from the inlet 54 to the nozzle 56. In comparison, theinlet 11 of the nacelle 10 shows significant increase in diameter fromthe cowl lip 13 to the maximum diameter of the nacelle 10 surroundingthe engine. Then, the diameter of the nacelle 10 reduces afterward alongthe nozzle 14 to the nozzle trailing edge 15. As would be understood bythose of skill in the art, a more streamlined nacelle design may produceweaker shocks and less overall drag.

The larger diameter nacelle 50 results in larger intake area for theinlet 54, consequently taking more air than is necessary or than theengine can handle. As a result, the bypass 58 may be used to capture theouter radial areas of the intake flow and bypass that flow around theengine. The ability to successfully bypass this flow around thepropulsion system may be enabled through the use of several additionaldesign features that facilitate the efficient capture, routing, andexhaust of the large quantity of bypass flow.

An embodiment of the invention may include a supersonic inlet forsupersonic aircraft that is configured to reduce the inlet'scontribution to a supersonic aircraft's sonic boom signature. Toaccomplish this, embodiments of the invention may position the cowl lipof the inlet such that the inlet captures the initial conic and/oroblique shock within the intake plane, preventing the conic shock energyor discontinuity from merging with the shocks generated by the airframeduring supersonic flight. It is also contemplated that the cowl angle ofthe nacelle may be reduced to zero or substantially zero in order toreduce the contribution of cowl shock and cowl drag on the overallsignature of a supersonic aircraft.

When designing an inlet in accordance with an embodiment of theinvention, a relaxed isentropic compression surface may be used. Asdiscussed in commonly owned U.S. patent application Ser. No. 11/639,339,filed Dec. 15, 2006 (entitled “Isentropic Compression Inlet forSupersonic Aircraft”), which is hereby incorporated by reference in itsentirety, a reduction in cowl angle may be achieved by designing aninlet to employ a relaxed isentropic compression surface such that thecowl angle may be reduced. A “relaxed isentropic compression” surface isan isentropic compression surface where a plurality of Mach lines do notfocus on the focus point where the initial shock and the terminal shockmeet. This lack of Mach line focusing may be configured to produce atotal level of compression less than the level of compression generatedby a conventional isentropic compression surface designed to the samecriteria. The relaxed isentropic compression surface may be configuredto increase terminal shock Mach number in the region of the cowl lip(creating the mechanism that reduces flow angle at the lip), but retainsa reasonable terminal shock Mach number along the remainder of theshock, including the base region of the terminal shock (preserving areasonable overall pressure recovery characteristic and good shockstability). Such an arrangement may significantly reduce the local flowangle at the cowl lip, leading to a reduction in cowling angle and asubstantial improvement in performance and a reduction in shockstrength.

FIG. 7A schematically illustrates a cross-section of the inlet 54. Asshown in FIG. 7A, the inlet 54 includes a cowl lip 60 and a leading edge64 of the external compression surface 66. The leading edge 64 generatesan initial shock 90. The compression surface 66 may be configured alongwith the leading edge 64 and the cowl lip 60 using a relaxed compressionarrangement to reduce the cowling angle of the cowl lip 60, and thestrength of the cowl shock 94, effectively reducing drag on the inlet 54as well as the inlet's contribution to the overall sonic boom signatureof the vehicle.

An interior splitter 68 functions within the inlet's subsonic diffuserto bifurcate the flow into a ‘primary’ stream 80 that enters the engine52 and a bypass stream 82 that circumvents the exterior of the enginethrough the bypass 58. As would be understood by those of skill in theart, the leading edge of the splitter 68 resides in a subsonic flowfield behind the terminal shock 92, allowing the leading edge of thesplitter 68 to use a blunted tip without detrimental performance impactat supersonic speeds.

FIG. 7B schematically illustrates a side view cross section of a relaxedisentropic external compression inlet 54 configured using shock-on-lipfocusing. The inlet 54 includes a compression surface 66 with an initialstraight surface 67 at an initial turn angle 66 a. The compressionsurface 66 includes a second compression surface 69 comprising a curvedsection 69 a and a straight section 69 b. The compression surface 66transitions into a shoulder 69 c, which defines the throat 71, thenarrowest portion of the inlet 54 flow path. The inlet 54 also includescowl lip 60 positioned at a cowl angle 60 b measured off the centerlineof the inlet 54. Although only the curved section 69 a of the secondcompression surface 69 generates isentropic compression, the entirecompression surface 66 is referred to herein as a relaxed isentropiccompression surface. For comparison, an example of a traditionalisentropic compression surface 66 a is shown in a dashed line. After theflow reaches throat 71, subsonic diffuser 73 provides a divergent flowpath delivering subsonic flow to the engine.

The inlet 54 first generates an initial shock 75 as the air flow inregion B travels in direction A and encounters the compression surface66 of inlet 54. The compression surface 66 may be configured to generatea terminal shock 77, having a base 77 a adjacent to the compressionsurface 66. As shown in FIG. 7B, the initial shock 75 and the terminalshock 77 are focused at a shock focus point 79. A cowl shock 81 is shownextending upward off the cowl lip 60. The relaxed isentropic compressionsurface allows for significant tailoring of the terminal shock 77 suchthat the outer radial region of the shock is nearly orthogonal to theinlet centerline. By shaping the terminal shock using relaxedcompression, the cowl lip 60 may be aligned with the local flow angle inthis outer radial region of the shock, greatly reducing the cowl lipangle. In addition, discrete adverse flow features, such as secondaryshock formation or flow separation, may be reduced at the cowl lipregion.

Although the cowl angle may be greatly reduced when using a relaxedisentropic compression inlet in accordance with FIG. 7B, the cowl lip isstill aligned with the local flow angle in the outer radial region ofthe terminal shock directly in front of the cowl lip. As would beunderstood by those of skill in the art, reducing the cowl angle 60 b,from the angle shown in FIG. 7B to zero or substantially zero may resultin flow distortion in the diffuser which may increase when the cowlingangle no longer aligns with the local flow in the vicinity of theterminal shock. This condition may generate secondary shocks and adversepressure fields in the vicinity of the cowl lip, which can introducestrong tip radial blockage defects in the flow seen by the engine at thefan face. Further, simply reducing the cowl angle 60 b to zero orsubstantially zero may also create temporal flow instability within thediffuser, potentially resulting from the flow disturbances created inthe outer radial region which may initiate and sustain diffuser flowresonance. Such resonance may adversely affect performance andpotentially damage the inlet and the engine.

Additionally, a simple reduction in cowl angle may be ineffective incontrolling aft cowling drag, or drag on the nacelle aft of the cowl lipresulting from any increase in nacelle diameter as the nacelle profileencompasses the engine. This increase in nacelle diameter may cause asharper gradient in the surface angle of the cowling as the maximumnacelle diameter is approached.

Furthermore, when the cowl lip is positioned to capture the initial orconic shock and the terminal shock in accordance with embodiments of theinvention, flow instabilities internal to the inlet may be introduced.As would be understood by those of skill in the art, the capture of theconic and terminal shocks may decrease the predictability of the postterminal shock flow environment and introduce flow separation on theinside cowl surface and produce unwanted flow dynamics.

As shown in FIG. 7C, the inlet structure and arrangement may beconfigured such that the cowl lip angle is extremely small or evenreduced to zero. As would be understood by those of skill in the art, azero or substantially zero cowl lip angle reduces the strength of thecowl shock due to reductions in the projected surface area exposed tothe freestream flow. Although the thickness of the cowl lip may includesome finite amount of material required to build the cowl lip, the cowllip structure may be extremely thin, depending on materials andapplication. It is contemplated that the nacelle wall thickness may growinward moving aft along the internal flowpath, providing the volumenecessary to incorporate structure while maintaining the uniformexternal diameter surface shape.

By employing a zero or substantially zero cowl lip angle, with referenceto a inlet axis 83, the region C may grow, especially if the nacelle isconfigured to fully encompass the engine without significant growth orcontraction in the outer diameter of the nacelle. Such a configurationmay reduce or eliminate the typical sharp growth of the outer diameterof the nacelle aft of the cowl lip as the nacelle encompasses theengine. As would be understood by those of skill in the art, a morecylindrical shape of uniform outer diameter may significantly reducecowling drag and cowl shock strength.

In accordance with embodiments of the invention, the nacelle bypass 58may be configured to handle the additional airflow that may enter theinlet due to the larger region C. By employing the bypass 58, the inlet54 may be configured to dispose of the excess flow, which wouldalternatively spill around the exterior of the cowl lip, creating higherdrag and defeating the objective of a lower sonic boom signature. Thenacelle bypass 58 avoids these spillage-related issues by routing theadditional flow through the nacelle and around the engine, eventuallyexhausting back to the free stream.

The nacelle bypass 58 may also serve to separate the flow distortioncaptured by the inlet 54. As discussed in U.S. patent application Ser.No. 11/639,339, the use of a relaxed isentropic compression surface 66may generate an initial shock 75 and a terminal shock 77, which may befocused at a point. The relaxed isentropic compression surface may alsobe configured to tailor the terminal shock 77 such that a region 85 ofrelaxed compression is produced. As a result, the strong velocitygradient in the outer radial region may generate the region 85 of flowdistortion. In accordance with embodiments of the invention, the bypass58 may be structured and arranged to separate the worst of the flowdistortion internal to the inlet 54 as shown as region 87. This region87 may include flow distortions introduced by the intersection of theinitial shock 75 and the terminal shock 77. In addition, the region 87may include flow distortion created by the sharp cowl lip 60, which mayproduce unfavorable flow distortion in the presence of cross-flow; forexample, when the vehicle experiences significant sideslip orangle-of-attack, or when the vehicle is subjected to high crosswindswhile operating on the ground.

More specifically, the bypass 58 operates to split the distorted flow inthe region 87 into the bypass 58, forming a bypass flow 82, which isseparated from the primary flow 80 by the splitter 68. The splitter 68prevents the bypass flow 82 and its inherent flow distortions fromreaching the sensitive turbomachinery. The resulting primary flow 80 maythen exhibit more uniform flow that may provide significant benefits toengine life and engine maintenance factors and improved fan andcompressor stability margins. The primary flow 80 profile may alsobenefit the engine performance by providing an increase in pressurerecovery that results from the removal of the more distorted, lowerpressure flow found in the region 87. The subsonic diffuser 73 may beconfigured to further slow the primary flow 80 into a subsonic flowsuitable for use by the engine. Also, the blunt leading edge 68 a ofbypass splitter 68 may be configured to couple favorably with cowl lip60 to produce a reduced flow distortion profile for the engine, similarto a traditional subsonic inlet.

The nacelle bypass 58 may also provide for the disposition of residualdiscrete flow defects or temporal flow instabilities, such as blockageprofiles resulting from flow separation or secondary shocks within thecowl lip area. The bypass 58 may work to eliminate resonance couplingbetween tip radial and centerbody boundary layer-related flow featuresthat can otherwise create adverse and strong instabilities, such asinlet buzz and other resonance types.

In accordance with embodiments of the invention, the inlet 54 maycapture the initial conic or oblique shock 75 within the intake plane ofinlet 54. Capturing the conic shock 75 may be accomplished by either aforward extension or movement of the cowling or by sizing the inlet to aMach number slightly lower than the design point. Although capturing theconic shock 75 would typically introduce large-scale flow instabilitiesfrom the interaction between the conic shock and the boundary layerimmediately aft of the cowl lip, the bypass 58 may be configured suchthat the conic shock 75 may be captured without significant impact onthe primary flow 80. As a result, the nacelle bypass 58 provides for aseparation, isolation, and disposal mechanism for the resulting spatialand temporal flow defects produced by conic shock capture, leaving theprimary flow path 80 significantly unaffected.

Referring back to FIG. 7A, the inlet 54 includes struts 70 and struts 72configured to stabilize the entire structure of the inlet 54. FIG. 8illustrates a plane view of the inlet 54 in FIG. 7A looking in thedirection of air flow at the face of the inlet 54.

As would be understood by those of skill in the art, the sharp leadingedge of the cowl lip 60 provides lower shock strength and dragcharacteristics at supersonic speeds compared to a configuration using amore blunt lip. However, sharp cowl lip inlet designs often produceunfavorable flow distortion in the presence of cross-flow, as when thevehicle is flying at significant sideslip or angle-of-attack or whensubjected to high cross-winds while operating on the ground. High flowdistortion within the diffuser subsequently enters the engine, reducingperformance and consuming engine operating stability margins. Byincluding the internal splitter and the bypass 58, the detrimentaleffects during low speed operation may be mitigated. As discussed inmore detail below, the blunt internal splitter leading edge 68 couplesfavorably with the sharp cowl lip 60 to produce a reduced flowdistortion profile for the engine face even at low speed or staticconditions. In effect, the sharp cowl lip 60 and the blunt leading edge68 function together to create a virtual inlet low speed inlet producinga low speed distortion flow, similar to a traditional subsonic inlet.

FIGS. 9A and 9B show perspective views of the inlet 54 with the outersurface of the inlet 54 drawn as transparent such that the internalstruts 72 of the bypass 58 may be clearly seen. FIG. 9C illustratesanother plane view of the inlet 54 in FIGS. 7A, 8, and 9A-B with theengine face shown at the back of the inlet 54. As shown in FIG. 9C, theengine fan 53 of engine 52 can be seen internal to the splitter leadingedge 68. The bypass 58 is shown external to the fan 53 of the engine 52such that the bypass flow 82 may flow around the engine 52.

Referring back to the blunt internal splitter leading edge 68, typicallow flight speeds or static operating conditions produce unfavorableflow characteristics for sharp cowl lips such as cowl lip 60. However,the blunt splitter leading edge 68 couples favorably with the sharp cowllip 60 to produce a reduced flow distortion profile for the engine faceeven at low speed or static conditions. FIG. 10A illustrates acomputational fluid dynamic (CFD) solution for a static or low speedcondition. As shown in FIG. 10A, the flow at low speed or staticconditions produces a large recirculating flow region directly under thecowl lip 60. As would be understood by those of skill in the art, such arecirculating flow condition would produce unfavorable engineperformance if the distortion aggravating effects of the recirculatingflow region reached the engine face. The blunt leading edge 68 producesa smoother flow tight against the inside surface of the internalsplitter. As shown in FIG. 10A, the smooth flow comes off the cowl lip60 and rides over the separated flow region before it encounters thesplitter leading edge 68. The leading edge 68 then traps the separatedflow between the two leading edges 60 and 68, allowing the flow to theengine face to exhibit a low distortion profile.

FIG. 10B illustrates this low distortion profile by plotting the axialcomponent of the velocity of the flow against the radial distance fromthe center of the nacelle at a location just forward of the engineentrance plane. The flow 80 in the primary path presents a generally lowdistortion profile as seen by the engine at low speed. As would beunderstood by those of skill in the art, without the internal splitterto separate the flow, the sharp edge of the cowl lip 60 would introducesignificant distortion to the engine face. The flow 82 in the bypasspath is represented with a negative velocity as air is pulled forwardthrough the bypass 58 from behind the engine. Again, as would beunderstood by those of skill in the art, engine stability margins andperformance increase correspondingly with reductions in distortion.Furthermore, the lower flow distortion may be used to eliminate thetypical requirement for low-speed distortion-attenuating auxiliaryintakes, reducing the complexity of the supersonic inlet 54.

FIG. 11A illustrates a CFD solution for an inlet in accordance withembodiments of the invention at a design speed of Mach 1.7. At highflight speed, the internal splitter may be configured to enhance theperformance benefit of isentropic relaxed compression inlet technology.The isentropic relaxed inlet compression permits a significant reductionin cowling angle, and associated drag and shock strength. But thisbenefit trades against a reduction in total pressure created by anunfavorable velocity gradient, generated by the compression surface 66,at the outer radial edges of the entrained flow. This velocity gradientproduces a strong flow distortion profile within the diffuser thatadversely impacts engine performance and stability margins as thedistorted flow is ingested by the engine. Although some engines may beconfigured to handle these radial velocity gradients, the positioning ofthe bypass splitter 68 may be configured to separate the outer radialflow, which contains the dramatic velocity gradient, from the primaryflow stream at high flight speed or supersonic speeds, preventing thedistorted flow from reaching and affecting the engine.

FIG. 11B illustrates the flow profile at Mach 1.7 design speed byplotting the axial component of the velocity of the flow against theradial distance from the center of the nacelle at a location justforward of the engine entrance plane. During supersonic flight, thesevere outer radial distortion pattern 82 produced by the compressionsurface may be confined primarily to the bypass path around the engine.The flow 80 follows the primary path to the engine face, illustrating agenerally smooth flow profile within the flow stream 80 entering theturbomachinery. As would be understood by those of skill in the art, theresult is an improvement in total pressure recovery at the engine faceand an increase in engine performance and stability margins.

FIG. 11C illustrates a Mach color computational fluid dynamics (CFD)solution for an inlet 11 in FIG. 1 employing a relaxed isentropiccompression design and shock-on-lip focusing with a cowl lip placed suchthat the conic shock is not captured by the inlet. FIG. 11 D illustratesa Mach color computational fluid dynamics (CFD) solution for an inlet 54in FIG. 5 in accordance with an embodiment of the invention. As withinlet 11, the inlet 54 employs a relaxed isentropic compression design.However, inlet 54 includes a near-zero cowl angle and is configured tocapture the conic shock internal to the inlet. FIGS. 11 C and 11 Drepresent inlets sized for a turbofan-type engine featuringapproximately 15,000 lb_(f) of maximum takeoff thrust and a moderatefan-to-compressor flow ratio of 3. Those areas of the flow fielddisturbed by less than 0.01 Mach number unit from the freestream Machnumber value are rendered white in both FIGS. 11C and 11D.

In comparison, the inlet 54 in FIG. 11D exhibits a greatly reduced shockdisturbance region 510 due to the zero-angle cowl and conic shockcapture. This may be easily seen by comparing the shock disturbanceregion 310 in FIG. 11C and the shock disturbance region 510 in FIG. 11D.In FIG. 11C, a large region 310 of disturbance is shown extending outand away from much of the forward nacelle surface. This indicates thatthe cowl shock 320, in FIG. 11C, is much stronger that the cowl shock520, in FIG. 11D. The strong cowl shock 320 will propagate away from thenacelle and eventually merge with shocks generated by aircraft airframe.In FIG. 11D, however, a relatively thin cowl shock disturbance 510extends out and away from only the very tip of the nacelle adjacent tothe zero-angle cowl lip. This is indicative of a much weaker cowl shock520 that will contribute less to the overall sonic boom signature.

Also illustrated in FIGS. 11C and 11D, the reduction in spillage may beseen for inlet 54 over inlet 11. As would be appreciated by one of skillin the art, the flow spillage 530 shown in FIG. 11D for the inlet 54 issignificantly less that the small amount of flow spillage 330 shown inFIG. 11C for the inlet 11. Specifically, FIG. 11D shows minimal spillageclose to the cowl lip, indicated by a significantly reduced cowl shockstrength. For inlet 54, these reductions in shock strength directlyreduce the inlet's contribution to a sonic boom signature for asupersonic aircraft employing inlet 54. As one of ordinary skill in theart will appreciate, the capture of the conic shock and the terminalshock 540 functions to virtually eliminate the flow spillage 530 and itsrelated contribution to shock strength. Moreover, the lack of anysignificant cowling profile (due to zero cowl angle) virtuallyeliminates cowl shock and cowl drag. The reduction in flow spillage 530also reduces drag.

FIG. 11D also illustrates a flow distortion 550 that is separated andisolated from the engine face. As discussed above, the zero orsubstantially zero cowl angle and the capture of the conic shock andterminal shock 540 may introduce flow distortions located in the outerradial region of the inlet. Although the bypass splitter is not shown inFIG. 11D, the flow distortion 550 adjacent to the cowl lip and the outersurface of the diffuser walls illustrates adverse flow characteristicsthat could be detrimental to the operability, performance, and life ofthe fan blades at an engine face. As discussed above, these adverse flowcharacteristics may be separated and isolated by the bypass 58.

The bypass may also provide significant attenuation of dynamic flowproperties produced by inlet design and operating characteristicstraditionally viewed as undesirable. For instance, the supersonic inletis typically constructed to position the initial conic or oblique inletshock outside of the cowl lip at the supersonic design point. Thisdesign technique results in increased flow spillage and drag, but isgenerally viewed as necessary to avoid unacceptable flow dynamics due toingestion of the initial shock. Such ingestion can produce a separatedflow region on the inside surface of the cowling that initiateshigh-amplitude flow oscillations detrimental to safe engine operationand potentially damaging to the structure of the inlet. By segregatingthe outer radial flow (that flow affected by directing the initial conicshock inside the cowl lip), the splitter 68 in FIG. 7A protects theprimary flow path to the engine face. As a consequence, the high-volumebypass stream may provide an independent zone, separated from theprimary flow path, that can serve to decouple and buffer the flowcommunication mechanics that otherwise drive ingested shock oscillation.

Referring back to FIGS. 7A, 8, and 9A-C, the bypass 58 itself can beemployed to handle structural load bearing duties typically assigned toan outer nacelle wall. In particular, support struts 72 in the bypass 58of the inlet 54 may be configured to couple the splitter 68 to the outerwall of the nacelle using a thin-wall composite structure, as anexample. Such an arrangement may be used to produce a stiff, strong, andlightweight nacelle structure, while maximizing the internal nacellevolume, which may then be used for bypass flow management. The struts inthe bypass stream can also be used to tailor the direction and amount ofairflow depending on local blockage characteristics within the bypassregion. For instance, the gearbox might completely block a significantcircumferential and radial portion of the bottom region of the bypassstream. The bypass struts in the inlet module can be used to bifurcatethe entrained bypass flow around the gearbox blockage. Another set ofstruts in the rear nacelle could be used to redirect the flow back intoa more circumferentially uniform pattern once aft of the gearbox.

FIG. 12 schematically illustrates a cross-sectional view of a nozzle 100with a bypass path 101 in accordance with an embodiment of theinvention. FIG. 13 illustrates a plane view of the nozzle 100 lookingupstream of the flow from aft of the nozzle. The nozzle 100 may beconfigured to exhaust the primary flow 102 from the engine and thebypass flow 104. The primary flow 102 and the bypass flow 104 areseparated by a bypass wall 106. The nozzle also includes a trailing edge108 on the bypass wall 106 and a trailing edge 110 on the outsidesurface of the nozzle 100. Struts 112 may be configured to couple theouter wall of the nozzle 100 with the bypass wall 106. As with thestruts in the inlet, the struts in the nozzle bypass flowpath may beconstructed from composite materials, as an example, and configured toreinforce the nozzle, producing a stiff, strong, and lightweightstructure.

FIGS. 14A and 14B show perspective views of the nozzle 100 with theouter surface of the nozzle 100 drawn as transparent such that theinternal struts 112 of the bypass 101 may be clearly seen. The outerwall of the nacelle may be constructed using thin-wall compositematerial construction techniques while using the size and placement ofthe struts for structural stiffening. Within the aft portion of thenozzle, as shown in FIGS. 13, 14A and 14B, the struts 112 may beconfigured with a variable thickness designed to control the exhaustexpansion of the flow 104 in the bypass 101. This may allow the bypass101 to be tailor-shaped in the aft portion of the nozzle bycircumferentially varying the strut 112 thickness, controlling theexhaust expansion requirements based on local flow conditions, such asin-flow pressure and mass flow, at the nozzle. As done for the primarystream 102, the exhaust flow expansion requirements for the bypass flow104 are determined by desired performance and sonic boomcharacteristics: the exhaust flow is usually fully expanded at designcondition to maximize thrust and to minimize the exhaust stream'sdisturbance of the external flow field, but it might also be designed topositively influence the performance and sonic boom characteristics ofthe primary exhaust stream 102. The bypass arrangement shown in FIGS.14A-B may maximize the use of linear surfaces for manufacturingsimplicity. For example, the struts 112 may be constructed using linearsurfaces while the thickness of the struts 112 may be tuned to producethe desired expansion characteristics based on local flow conditions.

FIG. 15 illustrates a CFD solution at freestream speed of Mach 1.7 ofthe internal flowpath and external flow region for a conventional nozzledesign as may be found on a high specific thrust military turbofanengine of approximately 11,000 lb_(f) maximum takeoff thrust class(non-afterburning). The traditionally configured nacelle and nozzle ofFIG. 15 produces a large nozzle boat tail angle due to the reduction inouter diameter as the nacelle approaches the trailing edge of thenozzle. The traditionally configured nacelle is shown in the figure toproduce an extensive external expansion fan and a subsequent strongre-shock due to the sharp and steep turning angles on the nozzle'sexterior surfaces.

FIG. 16 illustrates a CFD solution at freestream speed of Mach 1.7 ofthe internal flowpath and external flow region for the nozzle shown inFIG. 12 and consistent with a conventional turbofan type engine cyclefeaturing approximately 15,000 lbr of maximum takeoff thrust and amoderate fan-to-compressor flow ratio of 3. The bypass geometryillustrated in FIGS. 12, 13, and 14A-B, may be configured to permitimproved matching between nozzle and maximum nacelle diameter, allowingthe overall design of the nacelle to experience better streamlining ofthe outer nacelle and reduced nozzle boat tail angle and thereforeweaker expansion and re-shock regions. As a result, a nacelle inaccordance with embodiments of the invention may show an overallreduction in its contribution to the vehicle's sonic boom signature atsupersonic flight speeds while minimizing, or eliminating altogether,the drag impact resulting from the larger nacelle necessary toincorporate the bypass system.

FIG. 17 illustrates a Mach-based CFD solution for the flow 104 around anozzle bypass strut 112 shown in FIGS. 12, 13, and 14A-B for the Mach1.7 design condition. The solution image denotes a cut-through of awedged portion of the entire annular flow region, with a cross-sectionalcut of one bypass strut. The CFD analysis illustrates the concept ofcontrolling throat area through circumferential growth of the bypassstruts, providing stable choke line positioning and required mass flowand expansion characteristics. FIG. 17 indicates the location of thebypass nozzle choke line 120. The CFD analysis also indicates that thebypass struts 112, shown in FIGS. 12, 13, and 14A-B have negligibleimpact on exhaust flow characteristics. In addition to providingstructural performance, the thickness of the strut, shown in FIG. 17,may be used to control the expansion of the exhaust flow 104 through thebypass of the nozzle.

FIG. 18 schematically illustrates a 2 degree radial section of thenozzle 100 shown in FIGS. 12, 13, and 14A-B. Momentum-based and surfacepressure-based forces were extracted from viscous CFD solutions at Mach1.7 freestream consistent with the engine 52 shown in FIG. 5 operatingat design condition. The total axial force summation for thisconfiguration was determined to be 26,040 lb_(f), resulting in a netpropulsive force for the propulsion system adequate for meeting asupersonic vehicle cruise thrust requirement.

FIG. 19 illustrates another embodiment of the invention where a nacellebypass geometry 200 is shown in cylindrical coordinates. Additionalisometric and side views of nacelle bypass geometry 200 are also shownin FIGS. 20-27. As shown in FIG. 19, the nacelle bypass 200 isconfigured to direct the bypass flow around a gearbox fairing 202. Thenacelle bypass geometry 200 comprises an inlet portion 205 having a setof long aerodynamic vanes 210 and a set of short aerodynamic vanes 215.The combination of the long aerodynamic vanes 210 and the shortaerodynamic vanes 215 transforms an incoming airflow into a highsubsonic flow within the bypass. The long aerodynamic vanes 210 functionto direct air around blockages caused by the gearbox. As discussedabove, the aerodynamic vanes may be configured as struts. In FIG. 19,the long aerodynamic vanes 210 and the short aerodynamic vanes 215 mayalso function as struts between the inner and outer walls of the bypass.

As shown in FIG. 19, the airflow entering the inlet reaches the longaerodynamic vanes 210 and the short aerodynamic vanes 215 after passinga cowl highlight plane 220 and a splitter leading edge plane 225. As theairflow passes through the vanes 210 and 215, diffusion of the airflowreduces the speed of an incoming airflow. The airflow may be slowed froma supersonic flow at the cowl highlight plane 220 to a subsonic flow bythe time the airflow reaches the engine face plane 230. The flow remainssubsonic between a forward mount plane 235 and a choke plane 240 of anozzle section 245 of the nacelle bypass 200.

Nozzle section 245 of the nacelle bypass comprises a plurality of wedges250 formed by vanes. As with the vanes 210 and 215, the wedges 250 mayalso form structural struts to support the outer cowl and form thebypass structure. In another embodiment of the invention, the vanes 210and 215 may extend from the engine face plane 230, around the engine tomeet up with matching wedges 250, forming a plurality of independentbypass flow paths.

Once airflow reaches the choke plane 240, the wedges 250 may beconfigured to accelerate the airflow to supersonic speeds. In accordancewith one embodiment of the invention, the acceleration may beaccomplished by compressing the flow to the choke plane 240, where theflow would then go supersonic.

The wedges 250 control the airflow from the aft mount plane 255, theengine rear plane 260, a cowl trailing edge plane 265, and a nozzleshroud trailing edge 270. Supersonic acceleration may also be induced inthe airflow by configuring the wedges 250 such that an increase of thechannel cross-sectional area created by a continuous reduction in thesize of the wedges after the choke plane 255 and up to the trailing edge270. The shape of the wedges 250 and the increase of the channelcross-sectional area between the wedges may be optimized to allow forisentropic expansion of the airflow as it accelerates out of the rear ofthe nacelle bypass. The cowl may end at the cowl trailing edge plane 265with the wedges 250 extending beyond the cowl edge 265 and up to thenozzle shroud trailing edge 270. The wedges may split the airflow intoseparate regions as it exits the bypass which may help to prevent thecoalescence of shock waves produced by the bypass exhaust. Theprevention of the coalescence of shock waves may significantly reducethe strength of any shock wave formed off the trailing end of thenacelle, which is typically a strong contributor to the overall sonicboom characteristics of a supersonic vehicle.

FIG. 20 illustrates a front isometric view of a nacelle with the cowl280 shown as transparent so that the nacelle bypass geometry 200 (ofFIG. 19) is visible. In addition, the inlet 205, the compression surface275, and the splitter (or bypass wall) 203 are visible. The longaerodynamic vanes 210 and the short aerodynamic vanes 215 are disposedin the inlet section 205 between the bypass wall 203 and a cowl 280. Inaddition, the plurality of wedges 250 formed by vanes are also disposedin the nozzle section 245 between the bypass wall 203 and the cowl 280.

FIG. 21 illustrates an isometric rear view of the nacelle of FIG. 20with the cowl again shown in transparent. In FIG. 21, the nozzle 245 isshown with the nozzle centerbody 285 disposed internal to the bypasswall 203. From this view, the nozzle exit for the engine exhaust and theoutlet for the bypass flow can be seen together. It is also possible tosee how the wedges 250 extend beyond the trailing edge 265 of the cowl280.

FIG. 22 illustrates a side view of the nacelle of FIG. 20 with cowl 280shown in transparent. As illustrated, the aerodynamic vanes 210 and 215of the inlet section 205 and the wedges 250 of the nozzle section 245 donot connect with each other. As mentioned above, the long aerodynamicvanes 210 and the short aerodynamic vanes 215 may be configured suchthat they extend around the engine in a longitudinal direction in orderto connect with the matching wedges 250 of the nozzle section 245.

From the view shown in FIG. 22, a gearbox fairing 282 is created(internal to the cowl 280) by the vanes 210. The gearbox fairing 282provide necessary room to house the gearbox and other equipment used bythe engine. By creating this space for the gearbox, the outer surface ofthe cowl 280 may have limited or no protrusions or discontinuities,which may reduce interference and pressure drag and the nacelle'scontribution to the overall sonic boom signature of the vehicle.

FIG. 23 illustrates a front view of the nacelle of FIG. 20 and, inparticular, the inlet 205 of nacelle bypass geometry 200 (shown in FIG.19). As shown in the Figure, airflow that reaches the engine passesbetween the inlet compression surface 275 and internal to the bypasswall 203. Airflow bypassing the engine flows between the bypass wall 203and the cowl 280, where the long aerodynamic vanes 210 and the shortaerodynamic vanes 215 direct the airflow through the bypass and aroundthe gearbox fairing 282 (shown in FIG. 22). As discussed in reference toFIG. 19, the long aerodynamic vanes 210 and the short aerodynamic vanes215 may diffuse the incoming supersonic airflow to a subsonic flow. Astraight aerodynamic vane 290 runs along the top of the nacelle bypassgeometry 200. An aerodynamic vane 295 runs along the bottom of thenacelle bypass geometry 200. The vane 295 may be configured to bifurcatethe airflow, internal to the cowl 280, around the engine gearbox andother engine equipment.

FIG. 24 illustrates a rear view of the nacelle shown in FIG. 20 and, inparticular, the rear of the nozzle 245 of nacelle bypass geometry 200.Airflow exiting the engine passes between the nozzle centerbody 285 andthe bypass wall 203. Airflow that bypassed the engine exits by means ofchannels formed by wedges 250, the bypass wall 203, the cowl 280. Thewedges are tailored so that the bypass exhaust stream achieves a targetbypass exit Mach number, with the airflow that bypassed the engineexiting the nozzle section 245 at a supersonic speed. In at least oneembodiment of the invention, the target bypass exit Mach number for thebypass exhaust stream is approximately Mach 1.5.

FIG. 25 illustrates a viscous CFD solution of the airflow through thenacelle (including both the primary flow through the engine and thebypass flow through the nacelle bypass) shown in FIGS. 19-24. The Figureillustrates pressure coefficient data for the nacelle at Mach 1.7freestream. At inlet section 205, airflow is incident on the inletcompression surface 275. The airflow that is captured within the primaryflow path is directed to the engine. The remaining airflow captured bythe inlet passes through the nacelle bypass 200. As shown in the Figure,the airflow is compressed in inlet section 205, which reduces theairflow from supersonic speed to a subsonic airflow.

After passing around the gearbox fairing 282, the airflow reaches thenozzle section 245. The wedges 250 separate the flow into channels andexpands the flow into a supersonic exhaust flow around the nozzle. Asshown in the FIG. 25, the airflow undergoes expansion due to the shapeand configuration of the wedges 250, as explained above. As the airflowexpands back to ambient conditions, it experiences less pressure,resulting in the lowered pressure coefficient.

FIG. 26 is a close-up of the viscous CFD solution (shown in FIG. 25) ofthe airflow flowing only through the nacelle bypass. The CFD solution inFIG. 26 shows that the airflow is steered by the vanes 210 and 215 inthe inlet section of the nacelle. In addition, the CFD solution showsthat expansion of the airflow around the backside of gearbox fairing 282is accomplished without large-scale flow separation by the wedges 250.

FIG. 27 is a close-up of the viscous CFD solution (shown in FIG. 25) ofthe supersonic airflow entering the inlet. As described in detail above,the inlet compression surface 275 and the cowl lip or leading edge 315forms a terminal shock 310. The CFD solution shows that the terminalshock 310 is formed close to the cowl leading edge 315. This shows thatthe inlet effectively reduces airflow spillage outside the cowl leadingedge 315. As discussed in detail above, the low spillage indicates thatthe internal losses and blockage losses that occur by bypassing some ofthe airflow captured by the inlet have been effectively managed. Theseinternal losses and blockage issues would adversely affect the flowwithin the bypass. In addition, the CFD solution shows that the terminalshock 310 does not extend very far away from the cowl leading edge 315,which also indicates that spillage of airflow is being reduced and thenacelle bypass is capturing airflow that would ordinarily spill over ina conventional nacelle design.

It should be understood that a nacelle in accordance with an embodimentof the invention may be configured to use only the amount of bypass flownecessary to minimize shock characteristics to an acceptable level whilestill retaining adequate propulsion system performance. Such a designapproach to a nacelle configuration may balance various designcharacteristics, such as sonic boom requirements, design Mach number,vehicle size, propulsion integration requirements, engine type, missionperformance requirements, and mechanical complexity issues. Depending onthe application, it is contemplated that nacelle configurationsemploying bypass may even bypass a flow that exceeds that of the primaryflow. For example, a nacelle in accordance with embodiments of theinvention could bypass as much as 160 percent of the primary flow.

Bypass design may be used to minimize the total pressure losses alongthe length of the flowpath. In addition, some traditional engine mountdesigns, such as solid-web crane beams, may be redesigned and opened topermit pass-through of additional flow. The use of a hard-shell skinover heavily distributed regions of engine external hardware can also beused to reduce flow losses and protect sensitive external enginecomponents.

It also contemplated that the benefits of bypass may be maximizedthrough the use of thin-wall nacelle construction, trading conventionalstructural design techniques by employing the bypass struts 72 and 112,for example, as critical structural members. This technique provideslarger internal flowpath area, more diffusion potential, lower localvelocities, and less pressure loss. In fact, careful internal bypassdesign may avoid opportunities for local choking forward of the nozzlewhich would otherwise lead to excessive ram drag, poor nozzleperformance, and nondeterministic flow pumping characteristics.

The additional structural weight that would normally be incurred throughgrowing the diameter of a conventional nacelle can be minimized for thehigh-flow bypass concept through judicious use of composite material (toassist with the thin-wall construction technique), strut design andplacement, and reduced part-count due to reduced mechanical systemcomplexity (for example, eliminating auxiliary low speed intakesnormally used for distortion control).

Additionally, the bypass zone can also be utilized as a buffer regionbetween the outer nacelle wall and the engine surface. This hasimplications when integrating the nacelle with the airframe. Adverseinterference can be reduced when the outer wall shape along the lengthof the nacelle is tailored in a three-dimensional manner according tolocal flow characteristics near the airframe. This ability to tailor thewall shape is improved as the depth between the outer wall and theengine surface is increased, producing additional area and volume towork with. The bypass stream provides this additional depth along thelength of the flowpath, increasing the opportunities for localized,tailored, three-dimensional shaping of the nacelle surface.

The foregoing descriptions of specific embodiments of the invention asdefined by the claims below, including the preferred embodiments, arepresented for purposes of illustration and description. They are notintended to be exhaustive or to limit the invention to the precise formsdisclosed. Obviously, many modifications and variations are possible inview of the above teachings suited to particular uses.

1. A supersonic nozzle for a supersonic engine, comprising: an outerwall; a bypass wall disposed within the outer wall and configured toseparate an airflow inside the outer wall into a primary flow portionand a bypass flow portion, the primary flow portion passing through asupersonic engine and the bypass flow portion passing through a bypassthat bypasses the supersonic engine, the bypass flow portion comprisinga subsonic flow; and a set of struts configured to couple the outer wallwith the bypass wall, the set of struts tailoring a direction of thebypass flow portion, the set of struts further configured to direct thebypass flow portion around the supersonic engine and to expand thebypass flow portion into a supersonic exhaust, a first sub-set of theset of struts having a first curvature that directs the bypass flowportion in a first circumferential direction around a protruding portionof the supersonic engine, a second sub-set of the set of struts having asecond curvature that directs the bypass flow portion in a secondcircumferential direction around the protruding portion of thesupersonic engine, a third sub-set of the set of struts disposeddownstream of the first sub-set and having a third curvature thatdirects the bypass flow portion in the second circumferential direction,and a fourth sub-set of the set of struts disposed downstream of thesecond sub-set and having a fourth curvature that directs the bypassflow portion in the first circumferential direction.
 2. The supersonicnozzle of claim 1, wherein the set of struts comprises compositematerials for providing structural stiffness to the supersonic nozzle.3. The supersonic nozzle of claim 1, wherein a thickness of the strutsis configured to control an expansion of an exhaust from the supersonicengine.
 4. The supersonic nozzle of claim 1, wherein the set of strutsis constructed using linear surfaces.
 5. The supersonic nozzle of claim1, wherein the set of struts controls an amount of airflow depending onlocal blockage characteristics within the bypass flow portion; and theset of struts shapes the bypass flow portion around internal blockagescreated by a gearbox of the supersonic engine.
 6. The supersonic nozzleof claim 1, wherein the set of struts directs the bypass flow portioninto a substantially circumferentially uniform pattern prior to exhaust.7. The supersonic nozzle of claim 1, wherein the outer wall comprises atrailing edge that defines an exit cross-sectional area of thesupersonic nozzle.
 8. The supersonic nozzle of claim 1, wherein thesupersonic nozzle expands the bypass flow portion to maximize a thrustof the supersonic engine and to minimize a sonic boom signaturegenerated by exhaust of the primary flow portion.
 9. The supersonicnozzle of claim 1, wherein the outer wall comprises a thin-wallcomposite construction.
 10. The supersonic nozzle of claim 1, whereinthe set of struts receive the bypass flow portion and direct the bypassflow portion into a subsonic flow portion.
 11. A low shock supersonicnacelle, comprising: an engine; an outer wall; a bypass wall disposedwithin the outer wall and configured to support the engine; a set ofstruts configured to couple the outer wall with the bypass wall, a firstsub-set of the set of struts having a first curvature that directs thebypass flow portion in a first circumferential direction around aprotruding portion of the engine, a second sub-set of the set of strutshaving a second curvature that directs the bypass flow portion in asecond circumferential direction around the protruding portion of theengine, a third sub-set of the set of struts disposed downstream of thefirst sub-set and having a third curvature that directs the bypass flowportion in the second circumferential direction, and a fourth sub-set ofthe set of struts disposed downstream of the second sub-set and having afourth curvature that directs the bypass flow portion in the firstcircumferential direction; an inlet defined by front portions of theouter wall and the bypass wall, the inlet configured to decelerate anincoming airflow to a speed compatible with the engine; and a nozzledefined by rear portions of the outer wall and the bypass wall, thenozzle configured to accelerate an exhaust from the engine and bypass,wherein the bypass wall divides the incoming airflow into a primary flowportion directed into the engine and a bypass flow portion directed intoa bypass that bypasses the engine, the bypass flow portion comprising asubsonic flow, wherein the set of struts is further configured to directthe bypass flow portion around the engine and to expand the bypass flowportion into a supersonic exhaust.
 12. The low shock supersonic nacelleof claim 11, wherein the inlet comprises: a leading edge configured togenerate a first shock wave; a compression surface positioned downstreamof the leading edge and having at least one curved section configured togenerate compression; and a cowl lip on a cowling spatially separatedfrom the compression surface such that the cowl lip and the compressionsurface define an inlet opening for receiving a supersonic flow; whereinthe compression surface is configured to generate a second shock wavethat, during operation of the inlet at a predetermined cruise speed,extends from the compression surface to intersect the first shock waveat a point substantially adjacent to the cowl lip.
 13. The low shocksupersonic nacelle of claim 12, wherein the compression generated by thecurved section is characterized by a series of Mach lines where, duringoperation of the inlet at the predetermined cruise speed, at least aplurality of the Mach lines do not focus on the point substantiallyadjacent to the cowl lip.
 14. The low shock supersonic nacelle of claim11, wherein the bypass flow portion receives a diffused region of flowdistortion.
 15. The low shock supersonic nacelle of claim 11, furthercomprising a diffuser that receives the primary flow portion anddelivers a subsonic flow to the engine.
 16. The low shock supersonicnacelle of claim 11, wherein the set of struts comprises compositematerials configured to provide structural stiffness to the nozzle. 17.The low shock supersonic nacelle of claim 11, wherein a thickness of thestruts is configured to control an expansion of exhaust from the engine.18. The low shock supersonic nacelle of claim 11, wherein the set ofstruts is constructed using linear surfaces.
 19. The low shocksupersonic nacelle of claim 11, wherein the set of struts controls anamount of airflow depending on local blockage characteristics within thebypass; and the set of struts shapes the bypass flow portion aroundinternal blockages created by a gearbox of the engine.
 20. The low shocksupersonic nacelle of claim 11, wherein the set of struts directs thebypass flow portion into a substantially circumferentially uniformpattern prior to exhaust.
 21. The low shock supersonic nacelle of claim11, wherein the outer wall comprises a trailing edge that defines anexit cross-sectional area of the nozzle.
 22. The low shock supersonicnacelle of claim 11, wherein the nozzle expands the bypass flow portionto maximize a thrust of the supersonic engine and to minimize a sonicboom signature created by the primary flow portion.
 23. The low shocksupersonic nacelle of claim 11, wherein the outer wall comprises athin-wall composite construction.
 24. The low shock supersonic nacelleof claim 11, wherein increasing a distance between the outer wall andthe engine increases opportunities for localized, tailored,three-dimensional shaping of the outer wall.
 25. The low shocksupersonic nacelle of claim 11, wherein the bypass attenuatesinstabilities in the incoming airflow at the inlet.
 26. The low shocksupersonic nacelle of claim 11, wherein the set of struts receive thebypass flow portion and direct the bypass flow portion into a subsonicflow portion.
 27. A nacelle for a supersonic engine, comprising: anouter wall defining a closed volume; a bypass wall disposed within theclosed volume of the outer wall, the outer wall and the bypass wallcooperating to form a bypass that bypasses the supersonic engine; and aset of vanes configured to couple the outer wall with the bypass wall,defining a bypass therebetween, the set of vanes directing a bypassairflow into a subsonic flow, the set of vanes further configured todirect the subsonic flow around the supersonic engine and to expand thesubsonic flow into a supersonic exhaust, a first sub-set of the set ofvanes having a first curvature that directs the bypass flow portion in afirst circumferential direction around a protruding portion of thesupersonic engine, a second sub-set of the set of vanes having a secondcurvature that directs the bypass flow portion in a secondcircumferential direction around the protruding portion of thesupersonic engine, a third sub-set of the set of vanes disposeddownstream of the first sub-set and having a third curvature thatdirects the bypass flow portion in the second circumferential direction,and a fourth sub-set of the set of vanes disposed downstream of thesecond sub-set and having a fourth curvature that directs the bypassflow portion in the first circumferential direction.
 28. The nacelle ofclaim 27, wherein the set of vanes expands the subsonic flow into asubstantially circumferentially uniform pattern prior to exhaust. 29.The nacelle of claim 27, wherein the set of vanes controls an amount ofairflow depending on local blockage characteristics within the subsonicflow; and the set of vanes shapes the subsonic flow around internalblockages created by the gearbox of the supersonic engine.
 30. Thenacelle of claim 27, wherein the bypass wall is configured to separatean airflow inside the outer wall into a primary airflow and the bypassairflow, the primary airflow passing through a supersonic engine.
 31. Alow shock supersonic nacelle, comprising: an engine; an outer walldefining a closed volume; a bypass wall disposed within the closedvolume of the outer wall and configured to support the engine; a set ofvanes configured to couple the outer wall with the bypass wall, a firstsub-set of the set of vanes having a first curvature that directs thebypass flow portion in a first circumferential direction around aprotruding portion of the supersonic engine, a second sub-set of the setof vanes having a second curvature that directs the bypass flow portionin a second circumferential direction around the protruding portion ofthe supersonic engine, a third sub-set of the set of vanes disposeddownstream of the first sub-set and having a third curvature thatdirects the bypass flow portion in the second circumferential direction,and a fourth sub-set of the set of vanes disposed downstream of thesecond sub-set and having a fourth curvature that directs the bypassflow portion in the first circumferential direction; an inlet defined byfront portions of the outer wall and the bypass wall, the inletconfigured to decelerate an incoming airflow to a speed compatible withthe engine, the bypass wall dividing the incoming airflow into a primaryflow portion directed into the engine and a bypass flow portion directedinto a bypass that bypasses the engine, the bypass flow portioncomprising a subsonic flow; and a nozzle defined by rear portions of theouter wall and the bypass wall, the nozzle configured to accelerate anexhaust from the engine and the bypass, wherein the set of vanes isfurther configured to direct the bypass flow portion around the engineand to expand the bypass flow portion into a supersonic exhaust.
 32. Thelow shock supersonic nacelle of claim 31, wherein the inlet comprises: aleading edge configured to generate a first shock wave; a compressionsurface positioned downstream of the leading edge and having at leastone curved section configured to generate compression; and a cowl lip ona cowling spatially separated from the compression surface such that thecowl lip and the compression surface define an inlet opening forreceiving a supersonic flow; wherein the compression surface isconfigured to generate a second shock wave that, during operation of theinlet at a predetermined cruise speed, extends from the compressionsurface to intersect the first shock wave at a point substantiallyadjacent to the cowl lip.
 33. The low shock supersonic nacelle of claim31, further comprising: an inlet compression surface generating aterminal shock formed close to a leading edge of the outer wall.
 34. Thelow shock supersonic nacelle of claim 31, wherein the nozzle expands thebypass flow portion to maximize a thrust of the supersonic engine and tominimize a sonic boom signature created by the primary flow portion. 35.The low shock supersonic nacelle of claim 31, wherein the bypassattenuates instabilities in the incoming airflow at the inlet.